The present invention relates generally to land based gas turbines, for example, for electrical power generation, and more particularly to cooling the stage one nozzles of such turbines.
The traditional approach for cooling turbine blades and nozzles was to extract high pressure cooling air from a source, for example, from the intermediate and final stages of the turbine compressor. In such a system, a series of internal flow passages are typically used to achieve the desired mass flow objectives for cooling the turbine blades. In contrast, external piping is used to supply air to the nozzles, with air film cooling typically being used and the air exiting into the hot gas stream of the turbine. In advanced gas turbine designs, it has been recognized that the temperature of the hot gas flowing past the turbine components could be higher than the melting temperature of the metal. It was therefore necessary to establish a cooling scheme to protect the hot gas path components during operation. Steam has been demonstrated to be a preferred cooling media for cooling gas turbine nozzles (stator vanes), particularly for combined-cycle plants. See, for example, U.S. Pat. Ser. No. 5,253,976, the disclosure of which is incorporated herein by this reference. For a complete description of the steam-cooled buckets, reference is made to U.S. Pat. Ser. No. 5,536,143, the disclosure of which is incorporated herein by reference. For a complete description of the steam (or air) cooling circuit for supplying cooling medium to the first and second stage buckets through the rotor, reference is made to U.S. Pat. Ser. No. 5,593,274, the disclosure of which is incorporated herein by reference.
Because steam has a higher heat capacity than the combustion gas, however, it is considered inefficient to allow the coolant steam to mix with the hot gas stream. Consequently, in conventional steam cooled buckets it has been considered desirable to maintain cooling steam inside the hot gas path components in a closed circuit. Nevertheless, certain areas of the components in the hot gas path cannot practically be cooled with steam in a closed circuit. For example, the relatively thin structure of the trailing edges of the nozzle vanes effectively precludes steam cooling of those edges. Accordingly, air cooling is used to cool those portions of the nozzle vanes. For a complete description of the steam cooled nozzles with air cooling along the trailing edge, reference is made to U.S. Pat. Ser. No. 5,634,766, the disclosure of which is incorporated herein by reference.
In a typical closed loop steam or air cooled nozzle design such as that briefly described above and disclosed in the above-mentioned patents, the steam or air is used to cool the nozzle wall via impingement, or convection in the case of the trailing edge cavity. In some cases, with this kind of cooling scheme, the thermal gradient in the nozzle wall can reach very high levels, which can cause low LCF (Low Cycle Fatigue) life for local regions of the nozzle wall. Thus the inventors recognized that it would be desirable to modify the conventional closed loop cooled nozzle design to provide for a cooling of the exterior surface of the vane to reduce the local thermal gradient and in turn increase the local LCF life.
As noted above, since in a typical closed loop cooling circuit, the cooling media (steam or air) is at a pressure and/or temperature level different from that in the hot gas path, heretofore, by definition, such closed loop cooling circuits have excluded or isolated the closed loop cooling medium from the hot gas path. Indeed, heretofore it has been considered inefficient and undesirable for that cooling media to be introduced into the hot cooling path. The inventors have recognized, however, that by providing a small bleed of cooling media through suitably disposed openings in the airfoil wall of the otherwise closed loop cooling circuit, film cooling of the airfoil surface can be achieved to effectively increase the local LCF life in a manner that outweighs the potential efficiency loss. Thus, the invention is embodied in a vane or airfoil structure wherein a row or array of film cooling holes is defined to extend through the wall of the vane to communicate one or more of the interior nozzle cooling cavities with an exterior of the vane to allow a bleed flow of the cooling media through the nozzle airfoil wall to the hot gas path to form a cooling film to protect the airfoil. The film cooling holes are defined upstream of target low LCF life region(s) and can be disposed along a part or an entire radial length of the respective cavity, preferably corresponding to the location and extent of the local low LCF life region.
Thus, the present invention proposes to modify the typical closed loop steam or air cooled nozzle design by introducing cooling media, e.g. steam or air, film cooling to greatly reduce local thermal gradient, which, in turn, will increase the local LCF life. More specifically, the invention is embodied in the addition of at least one film cooling hole, and more preferably an array of film cooling holes to a closed loop steam or air cooled nozzle for providing a cooling media source for film cooling of the airfoil surface in regions where low LCF life would otherwise exist due to high thermal gradient. The film cooling holes are defined through the wall of one or more cavities of a closed loop steam or air cooled gas turbine nozzle. Cooling media with thus flow out into the hot gas path through film holes.
Accordingly, in an embodiment of the present invention, there is provided a cooling system for cooling the hot gas components of a nozzle stage of a gas turbine, in which closed circuit steam or air cooling and/or open circuit air cooling systems may be employed. In the closed circuit system, a plurality of nozzle vane segments are provided, each of which comprises one or more nozzle vanes extending between radially inner and outer walls. The vanes have a plurality of cavities in communication with compartments in the outer and inner walls for flowing cooling media in a closed circuit for cooling the outer and inner walls and the vanes per se. This closed circuit cooling system is substantially structurally similar to the steam cooling system described and illustrated in the prior referenced U.S. Pat. Ser. No. 5,634,766, with certain exceptions as noted below. Thus, cooling media may be provided to a plenum in the outer wall of the segment for distribution to chambers therein and passage through impingement openings in a plate for impingement cooling of the outer wall surface of the segment. The spent impingement cooling media flows into leading edge and aft cavities extending radially through the vane. At least one cooling fluid return/intermediate cooling cavity extends radially and lies between the leading edge and aft cavities. A separate trailing edge cavity may also provided. The flow of cooling air in a trailing edge cavity per se is the subject of a U.S. Pat. Ser. No. 5,611,662, the disclosure of which is incorporated herein by reference. The cooling air from that trailing edge cavity flows to the inner wall, for flow through a passage for supplying purge air to the wheelspace, or into the hot gas path. To cool the airfoil surface in regions where low LCF life will otherwise exist due to high thermal gradient, at least one film cooling hole is defined through the wall of one or more of the aforementioned cavities of the closed loop steam or air cooled gas turbine nozzle. Cooling media then flows out into the hot gas path through film cooling hole(s) defined in the airfoil wall, thereby to create a cooling film to cool the airfoil surface.
More specifically, in a preferred embodiment of the present invention, there is provided a closed circuit stator vane segment comprising radially inner and outer walls spaced from one another, a vane extending between the inner and outer walls and having leading and trailing edges, the vane including discrete leading edge, trailing edge and intermediate cavities between the leading and trailing edges and extending radially of the vane, said leading edge and intermediate cavities together defining a substantially closed cooling circuit for flow of cooling media through said vane, an insert in the leading edge cavity for receiving cooling media and having impingement openings for directing the cooling media against interior wall surfaces of the leading edge cavity for impingement cooling of the vane about the leading edge cavity, an insert in the intermediate cavity for receiving cooling media and having impingement openings for directing the cooling media against interior wall surfaces of the intermediate cavity for impingement cooling of the vane about the intermediate cavity, the trailing edge cavity lying in communication with a cooling air inlet for receiving cooling air therefrom and having an outlet one of at a trailing edge thereof and at a radially inner end thereof, for directing spent cooling air one of into the hot gas path exterior to the vane and into a wheelspace between adjacent turbine stages, and wherein at least one film cooling hole is defined through a wall of at least one of the cavities for flow communication between an interior of the vane cavity and an exterior of the vane, to cool the airfoil surface and thus reduce the thermal gradient in that region.
The present invention may further be embodied in a substantially closed circuit cooling system for cooling the hot gas components of nozzle stages of a gas turbine, particularly the first nozzle stage, modified to provide for film cooling for certain of those components. More particularly, nozzle vane segments are provided having the necessary structural integrity under high thermal fluxes and pressures affording a capacity of being cooled by a cooling medium, preferably steam, flowing in a pressurized substantially closed circuit. Thus, the present invention provides, in at least the first stage of a turbine, a plurality of nozzle vane segments each of which comprise one or more nozzle vanes extending between radially outer and inner walls. The vanes have a plurality of cavities in communication with compartments in the outer and inner walls for flowing a cooling media, preferably steam, in a substantially closed-circuit path for cooling the outer and inner walls and the vanes, per se. Impingement cooling is provided in the leading cavity of the vane, as well as in the intermediate, return cavity(ies) of the first stage nozzle vane. Inserts in the leading and aft cavities comprise sleeves that extend through the cavities spaced from the walls thereof. The inserts have impingement holes in opposition to the walls of the cavity whereby steam flowing into the inserts flows outwardly through the impingement holes for impingement cooling of the vane walls. Return channels are provided along the inserts for channeling the spent impingement cooling steam. Similarly, inserts in the return, intermediate cavity(ies) have impingement openings for flowing impingement cooling medium against the side walls of the vane. Those inserts also have return cavities for collecting the spent impingement cooling steam and transmitting it to the cooling medium, e.g. steam, outlet.
The first stage nozzle segments further provide for film cooling of the airfoil surface in regions where low LCF life will otherwise exist due to high thermal gradient. More particularly, at least one film cooling hole and preferably a plurality of or an array of film cooling holes are defined in or along at least a portion of the wall of at least one cavity of the segment for bleeding a portion of the cooling medium from the otherwise closed circuit to film cool a predetermined portion of the vane exterior.